Development of high burn rate propellant and testing in miniature rocket motor for control applications

Abstract

For aerospace vehicles, notably missiles or satellites for attitude and divert control, impulsive thrusters are most desirable if they are quantized as exact instantaneous pulses available over and after extended periods of inactive storage. These thrusters operate for a few milliseconds (ms), typically between 10 ms to 100 ms duration. Compared to its liquid or gaseous counterpart, an impulsive thruster powered by solid propellants is a simple, and reliable solution. Therefore, they are suitable as reaction control systems and divert thrusters for satellite and unmanned aerial vehicle applications.

A typical solid propellant would need an extremely thin web thickness for port burning arrangement due to its brief burn duration, which makes propellant manufacturing and its structural anchoring in motor challenging. A high burn rate and high-density propellant based on ammonium perchlorate, aluminum, and Teflon is processed to enable end-burning configuration thus solving the problems, simplifying propellant manufacturing, structural design and integration. It also permits stacking propellants for multi-pulse applications. The high burn rate (40 mm/s at 120 bar pressure) propellant is formulated and tested for ballistic performance evaluation in proof motor configuration, and evaluation of its advantages over conventional propellant. A case of identical design requirements with both double base propellant and high burn rate pressed propellant is compared, tested and evaluated to compare the relative performance merits. The volumetric load is increased from 47% to 82% due to the development and adoption of high burn rate propellant. The overall weight of the thruster is also reduced by 20%.